EP2864707A1 - Turbine engine combustor wall with non-uniform distribution of effusion apertures - Google Patents
Turbine engine combustor wall with non-uniform distribution of effusion aperturesInfo
- Publication number
- EP2864707A1 EP2864707A1 EP20130807403 EP13807403A EP2864707A1 EP 2864707 A1 EP2864707 A1 EP 2864707A1 EP 20130807403 EP20130807403 EP 20130807403 EP 13807403 A EP13807403 A EP 13807403A EP 2864707 A1 EP2864707 A1 EP 2864707A1
- Authority
- EP
- European Patent Office
- Prior art keywords
- apertures
- effusion
- axial region
- heat shield
- impingement
- Prior art date
- Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
- Granted
Links
- 238000009828 non-uniform distribution Methods 0.000 title description 3
- 238000010791 quenching Methods 0.000 claims abstract description 51
- 238000002485 combustion reaction Methods 0.000 claims description 35
- 238000011144 upstream manufacturing Methods 0.000 claims description 24
- 238000001816 cooling Methods 0.000 description 21
- 239000000446 fuel Substances 0.000 description 9
- 230000000712 assembly Effects 0.000 description 4
- 238000000429 assembly Methods 0.000 description 4
- 239000000463 material Substances 0.000 description 3
- 239000012720 thermal barrier coating Substances 0.000 description 3
- 239000000919 ceramic Substances 0.000 description 2
- 239000000203 mixture Substances 0.000 description 2
- 230000004888 barrier function Effects 0.000 description 1
- RLQJEEJISHYWON-UHFFFAOYSA-N flonicamid Chemical compound FC(F)(F)C1=CC=NC=C1C(=O)NCC#N RLQJEEJISHYWON-UHFFFAOYSA-N 0.000 description 1
Classifications
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
- F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
- F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
- F23R3/02—Continuous combustion chambers using liquid or gaseous fuel characterised by the air-flow or gas-flow configuration
- F23R3/04—Air inlet arrangements
- F23R3/06—Arrangement of apertures along the flame tube
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
- F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
- F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
- F23R3/002—Wall structures
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
- F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
- F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
- F23R3/02—Continuous combustion chambers using liquid or gaseous fuel characterised by the air-flow or gas-flow configuration
- F23R3/04—Air inlet arrangements
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
- F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
- F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
- F23R3/02—Continuous combustion chambers using liquid or gaseous fuel characterised by the air-flow or gas-flow configuration
- F23R3/04—Air inlet arrangements
- F23R3/10—Air inlet arrangements for primary air
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
- F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
- F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
- F23R3/42—Continuous combustion chambers using liquid or gaseous fuel characterised by the arrangement or form of the flame tubes or combustion chambers
- F23R3/50—Combustion chambers comprising an annular flame tube within an annular casing
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05B—INDEXING SCHEME RELATING TO WIND, SPRING, WEIGHT, INERTIA OR LIKE MOTORS, TO MACHINES OR ENGINES FOR LIQUIDS COVERED BY SUBCLASSES F03B, F03D AND F03G
- F05B2260/00—Function
- F05B2260/20—Heat transfer, e.g. cooling
- F05B2260/201—Heat transfer, e.g. cooling by impingement of a fluid
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05B—INDEXING SCHEME RELATING TO WIND, SPRING, WEIGHT, INERTIA OR LIKE MOTORS, TO MACHINES OR ENGINES FOR LIQUIDS COVERED BY SUBCLASSES F03B, F03D AND F03G
- F05B2260/00—Function
- F05B2260/20—Heat transfer, e.g. cooling
- F05B2260/202—Heat transfer, e.g. cooling by film cooling
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
- F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
- F23R2900/00—Special features of, or arrangements for continuous combustion chambers; Combustion processes therefor
- F23R2900/03041—Effusion cooled combustion chamber walls or domes
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
- F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
- F23R2900/00—Special features of, or arrangements for continuous combustion chambers; Combustion processes therefor
- F23R2900/03044—Impingement cooled combustion chamber walls or subassemblies
Definitions
- This disclosure relates generally to a turbine engine combustor and, more particularly, to a turbine engine combustor wall with a non-uniform distribution of effusion apertures.
- a combustor wall for a turbine engine with an axial centerline.
- the combustor wall includes a combustor support shell and a combustor heat shield.
- the support shell includes a plurality of shell quench apertures, a plurality of first impingement apertures, and a plurality of second impingement apertures.
- the heat shield includes a plurality of shield quench apertures fluidly coupled with the shell quench apertures, a plurality of first effusion apertures fluidly coupled with the first impingement apertures, and a plurality of second effusion apertures fluidly coupled with the second impingement apertures.
- the shield quench apertures and the first effusion apertures are configured in a first axial region of the heat shield.
- the second effusion apertures are configured in a second axial region of the heat shield located axially between the first axial region and a downstream end of the heat shield.
- a density of the first effusion apertures in the first axial region is greater than a density of the second effusion apertures in the second axial region.
- an axial flow combustor for a turbine engine with an axial centerline.
- the combustor includes a first combustor wall, a second combustor wall with a support shell and a heat shield, and an annular combustor bulkhead extending radially between an upstream end of the first combustor wall and an upstream end of the second combustor wall.
- the support shell includes a plurality of shell quench apertures, a plurality of first impingement apertures, and a plurality of second impingement apertures.
- the heat shield includes a plurality of shield quench apertures fluidly coupled with the shell quench apertures, a plurality of first effusion apertures fluidly coupled with the first impingement apertures, and a plurality of second effusion apertures fluidly coupled with the second impingement apertures.
- the shield quench apertures and the first effusion apertures are configured in a first axial region of the heat shield.
- the second effusion apertures are configured in a second axial region of the heat shield.
- the first axial region is located axially between the upstream end of the second combustor wall and the second axial region.
- a density of the first effusion apertures in the first axial region is greater than a density of the second effusion apertures in the second axial region.
- the first combustor wall may be disposed radially within the second combustor wall.
- the second combustor wall may be disposed radially within the first combustor wall.
- the support shell also includes a plurality of third impingement apertures
- the heat shield also includes a plurality of third effusion apertures, which are fluidly coupled with the third impingement apertures.
- the third effusion apertures are configured in a third axial region of the heat shield located axially between the first axial region and an upstream end of the heat shield. A density of the third effusion apertures in the third axial region is less than the density of the first effusion apertures in the first axial region.
- the density of the third effusion apertures in the third axial region is greater than the density of the second effusion apertures in the second axial region.
- the support shell also includes a plurality of third impingement apertures
- the heat shield also includes a plurality of third effusion apertures, which are fluidly coupled with the third impingement apertures.
- Axes of more than seventy five percent of the third effusion apertures extend circumferentially through the panel and are substantially perpendicular to the axial centerline.
- the third effusion apertures are configured in a third axial region of the heat shield located axially between the first axial region and an upstream end of the heat shield.
- a density of the third effusion apertures in the third axial region may be substantially equal to the density of the first effusion apertures in the first axial region.
- a plurality of the first effusion apertures located adjacent to a first of the panel quench apertures, have axes that are substantially tangent to a downstream side of the first panel quench aperture.
- the impingement apertures are configured to exhibit a pressure drop across the support shell, and the effusion apertures are configured to exhibit a pressure drop across the heat shield.
- a ratio of the pressure drop across the support shell to the pressure drop across the heat shield can be between about 2:1 and about 9:1.
- some or all of the impingement apertures and some or all of the effusion apertures have substantially equal diameters. In other embodiments, the diameters of some or all of the effusion apertures are greater than diameters of some or all of the impingement apertures. In still other embodiments, the diameters of some or all of the effusion apertures are less than diameters of some or all of the impingement apertures.
- axes of some or all of the effusion apertures are offset from a combustion chamber surface of the heat shield by between about fifteen and about thirty degrees, and/or axes of some or all of the impingement apertures are substantially perpendicular to an impingement cavity surface of the support shell.
- an impingement cavity extends radially between the support shell and the heat shield, and fluidly couples some or all of the impingement apertures with some or all of the effusion apertures.
- the support shell has an annular cross-sectional geometry and extends axially between an upstream end and a downstream end.
- the heat shield has an annular cross-sectional geometry and extends axially between an upstream end and the downstream end of the panel.
- the heat shield is disposed radially within the support shell.
- the support shell is disposed radially within the heat shield.
- the heat shield includes a plurality of circumferential heat shield panels and/or a plurality of axial heat shield panels.
- the first axial region and/or the second axial region includes a plurality of circumferential first sub-regions and a plurality of circumferential second sub-regions.
- a density of the effusion apertures in each first sub-region is greater than a density of the effusion apertures in each second sub-region.
- the density of the effusion apertures in the respective axial region is equal to an average or mean of the densities of the effusion apertures in the first sub-regions and the densities of the effusion apertures in the second sub-regions.
- the shell quench apertures and the first impingement apertures are configured in a first axial region of the support shell, and the second impingement apertures are configured in a second axial region of the support shell located axially between the first axial region of the support shell and a downstream end of the support shell.
- a density of the first impingement apertures in the first axial region of the support shell is greater than a density of the second impingement apertures in the second axial region of the support shell.
- FIG. 1 is a side- sectional diagrammatic illustration of a turbine engine combustor.
- FIG. 2 is a cross-sectional diagrammatic illustration of a turbine engine combustor.
- FIG. 3 is an exploded, perspective diagrammatic illustration of a section of a combustor wall.
- FIG. 5 is a diagrammatic illustration of a section of a combustor heat shield.
- FIG. 6 is a side-sectional diagrammatic illustration of a combustor wall.
- the first combustor wall 16 and the second combustor wall 20 can each include a combustor support shell 30 and a combustor heat shield 32.
- the support shell 30 extends axially between the upstream end 14, 18 and a downstream end 34, 36.
- the support shell 30 extends circumferentially around the axial centerline 24, which provides the support shell 30 with an annular cross-sectional geometry. Referring to FIG. 3, the support shell 30 also extends radially between a combustor plenum surface 38 and a first impingement cavity surface 40.
- the support shell 30 can be constructed as a single integral tubular body. Alternatively, the support shell 30 can be assembled from a plurality of circumferential support shell panels and/or a plurality of axial support shell panels.
- the support shell 30 includes a plurality of shell quench apertures 42 and a plurality of impingement apertures (e.g., the apertures 44).
- the shell quench apertures 42 extend radially through the support shell 30 between the combustor plenum surface 38 and the first impingement cavity surface 40.
- Each of the shell quench apertures 42 can have a circular cross-sectional geometry with a first diameter 46.
- the impingement apertures (e.g., the apertures 44) extend radially through the support shell 30 between the combustor plenum surface 38 and the first impingement cavity surface 40.
- Each of the impingement apertures (e.g., the apertures 44) has an axis 48 that is angularly offset from first impingement cavity surface 40, for example, by an angle ⁇ of about ninety degrees.
- Each of the impingement apertures (e.g., the apertures 44) can have a circular cross-sectional geometry with a second diameter 50, which is substantially (e.g., at least five to twenty times) smaller than the first diameter 46.
- the impingement apertures can include a plurality of first impingement apertures 52, a plurality of second impingement apertures 54, a plurality of third impingement apertures 56, a plurality of fourth impingement apertures 44, a plurality of fifth impingement apertures 58, and a plurality of sixth impingement apertures 60.
- the shell quench apertures 42 and the impingement apertures can be arranged in one or more support shell cooling regions.
- the first impingement apertures 52 for example, are arranged in a first axial region 62.
- the first axial region 62 extends axially from a second axial region 64 towards the upstream end 14, 18, and circumferentially around the centerline 24.
- the second impingement apertures 54 are arranged in the second axial region 64.
- the second axial region 64 extends axially from the first axial region 62 to a third axial region 66, and circumferentially around the centerline 24.
- the third impingement apertures 56 are arranged in the third axial region 66.
- the third axial region 66 extends axially from the second axial region 64 to a fourth axial region 68, and circumferentially around the centerline 24.
- the shell quench apertures 42 and the fourth impingement apertures 44 are arranged in the fourth axial region 68.
- the fourth axial region 68 extends axially from the third axial region 66 to a fifth axial region 70, and circumferentially around the centerline 24.
- the fifth impingement apertures 58 are arranged in the fifth axial region 70.
- the fifth axial region 70 extends axially from the fourth axial region 68 to a sixth axial region 72, and circumferentially around the centerline 24.
- the sixth impingement apertures 60 are arranged in the sixth axial region 72.
- the sixth axial region 72 extends axially from the fifth axial region 70 towards (e.g., to) the downstream end 34, 36, and circumferentially around the centerline 24.
- the number of and relative spacing between the impingement apertures included in each of the support shell cooling regions is selected to provide each cooling region with a respective impingement aperture density.
- the term "impingement aperture density" describes a ratio of the number of impingement apertures included in a unit (e.g., a square inch) of substantially unobstructed support shell surface area.
- Unobstructed support shell surface area can include, for example, portions of the first impingement cavity surface 40 that do not include non-cooling apertures (e.g., the shell quench apertures 42) and/or other support shell features such as, for example, bosses, studs, flanges, rails, etc. connected to the combustor plenum surface 38.
- Obstructed support shell surfaces can include, for example, first regions 74 of the first impingement cavity surface opposite shell quench aperture 42 rails, and second regions 76 of the first impingement cavity surface opposite stud apertures.
- the support shell 30 includes !Nf number of the first impingement apertures 52, which provides the first axial region 62 with a first impingement aperture density.
- the support shell 30 includes N 2 number of the second impingement apertures 54, which provides the second axial region 64 with a second
- the fifth impingement aperture density is, for example, less than the second, third and fourth impingement aperture densities, and substantially equal to the first impingement aperture density.
- the support shell 30 includes N 6 number of the sixth impingement apertures 60, which provides the sixth axial region 72 with a sixth impingement aperture density.
- the sixth impingement aperture density is, for example, greater than the fifth impingement aperture density, and substantially equal to or less than the fourth impingement aperture density.
- the impingement aperture density in one or more of the support shell cooling regions may change (e.g., intermittently increase and decrease) as the region extends circumferentially around the centerline 24.
- the second axial region 64 includes a plurality of (e.g., triangular, trapezoidal, etc.) circumferential first sub-regions 78 and a plurality of (e.g., triangular, trapezoidal, etc.) circumferential second sub-regions 80.
- the first sub-regions 78 are configured to be
- the heat shield 32 extends axially between an upstream end 82 and a downstream end 84.
- the heat shield 32 extends circumferentially around the axial centerline 24, which provides the heat shield 32 with an annular cross-sectional geometry.
- the heat shield 32 also extends radially between a second impingement cavity surface 86 and a combustion chamber surface 88.
- the heat shield 32 can be assembled from a plurality of circumferential heat shield panels 90 and 92 and/or a plurality of axial heat shield panels 90 and 92.
- the heat shield 32 can be constructed as a single integral tubular body.
- the heat shield 32 includes a plurality of shield quench apertures 94 and a plurality of effusion apertures (e.g., the apertures 96).
- the shield quench apertures 94 extend radially through the heat shield 32 between the second impingement cavity surface 86 and the combustion chamber surface 88.
- Each of the shield quench apertures 94 can have a circular cross-sectional geometry with a third diameter 98.
- the third diameter 98 may be less than the first diameter 46 where, for example, the heat shield 32 includes annular flanges that nest within the shell quench apertures 42 and fluidly couple the shield quench apertures 94 to the shell quench apertures 42.
- the third diameter 98 may be greater than or equal to the first diameter 46.
- the effusion apertures extend radially through the heat shield 32 between the second impingement cavity surface 86 and the combustion chamber surface 88.
- Each of the effusion apertures (e.g., the apertures 96) has an axis 100 that is angularly offset from the combustion chamber surface 88, for example, by an angle a of between about fifteen and about thirty degrees (e.g., about 25°).
- Each of the effusion apertures (e.g., the apertures 96) can have a circular cross-sectional geometry with a fourth diameter 102, which is substantially (e.g., at least five to twenty times) smaller than the third diameter 98.
- the fourth diameter 102 of some or all of the effusion apertures can be greater than, less than or equal to the second diameter 50.
- the effusion apertures can include a plurality of first effusion apertures 104, a plurality of second effusion apertures 106, a plurality of third effusion apertures 108, a plurality of fourth effusion apertures 96, a plurality of fifth effusion apertures 110, and a plurality of sixth effusion apertures 112.
- the shield quench apertures 94 and the effusion apertures can be arranged in one or more heat shield cooling regions.
- the first effusion apertures 104 for example, are arranged in a first axial region 114.
- the first axial region 114 extends axially from a second axial region 116 towards (e.g., to) the upstream end 82, and circumferentially around the centerline 24.
- the second effusion apertures 106 are arranged in the second axial region 116.
- the second axial region 116 extends axially from the first axial region 114 to a third axial region 118, and circumferentially around the centerline 24.
- the third effusion apertures 108 are arranged in the third axial region 118.
- the third axial region 118 extends axially from the second axial region 116 to a fourth axial region 120, and circumferentially around the centerline 24.
- the shield quench apertures 94 and the fourth effusion apertures 96 are arranged in the fourth axial region 120.
- the fourth axial region 120 extends axially from the third axial region 118 to a fifth axial region 122, and circumferentially around the centerline 24.
- the fifth effusion apertures 110 are arranged in the fifth axial region 122.
- the fifth axial region 122 extends axially from the fourth axial region 120 to a sixth axial region 124, and circumferentially around the centerline 24.
- the sixth effusion apertures 112 are arranged in the sixth axial region 124.
- the sixth axial region 124 extends axially from the fifth axial region 122 towards (e.g., to) the downstream end 84, and circumferentially around the centerline 24.
- the number of and relative spacing between the effusion apertures included in each of the heat shield cooling regions is selected to provide each cooling region with a respective effusion aperture density.
- the term "effusion aperture density" describes a ratio of the number of effusion apertures included in a unit (e.g., a square inch) of substantially unobstructed heat shield surface area.
- Unobstructed heat shield surface area can include, for example, portions of the combustion chamber surface 88 that do not include non-cooling apertures (e.g., the shield quench apertures 94) and/or other heat shield features such as, for example, bosses, studs, flanges, rails, etc. connected to the second impingement cavity surface 86.
- Obstructed heat shield surfaces can include, for example, first regions 128 of the combustion chamber surface opposite shell quench aperture 94 rails, and second regions 130 of the combustion chamber surface opposite studs.
- the heat shield 32 includes M 1 number of the first effusion apertures 104, which provides the first axial region 114 with a first effusion aperture density.
- the heat shield 32 includes M 2 number of the second effusion apertures 106, which provides the second axial region 116 with a second effusion aperture density that is, for example, greater than the first effusion aperture density.
- the heat shield 32 includes M 3 number of the third effusion apertures 108, which provides the third axial region 118 with a third effusion aperture density that is, for example, greater than (or substantially equal) to the second effusion aperture density.
- the heat shield 32 includes M 4 number of the fourth effusion apertures 96, which provides the fourth axial region 120 with a fourth effusion aperture density that is, for example, substantially equal to the third effusion aperture density.
- the heat shield 32 includes M 5 number of the fifth effusion apertures 110, which provides the fifth axial region 122 with a fifth effusion aperture density.
- the fifth effusion aperture density is, for example, less than the second, third and fourth effusion aperture densities, and substantially equal to the first effusion aperture density.
- the heat shield 32 includes M 6 number of the sixth effusion apertures 112, which provides the sixth axial region 124 with a sixth effusion aperture density.
- the sixth effusion aperture density is, for example, greater than the fifth effusion aperture density, and substantially equal to or less than the fourth effusion aperture density.
- the effusion aperture density in one or more of the heat shield cooling regions may change (e.g., intermittently increase and decrease) as the region extends circumferentially around the centerline 24.
- the second axial region 116 includes a plurality of (e.g., triangular, trapezoidal, etc.) circumferential first sub-regions 132 and a plurality of (e.g., triangular, trapezoidal, etc.) circumferential second sub-regions 134.
- the first sub-regions 132 are configured to be circumferentially aligned with the fuel injector assemblies 22.
- Each of the second sub-regions 134 extends circumferentially between two respective first sub-regions 132.
- the density of the second effusion apertures 106 in the first sub-regions 132 is greater than that of the second sub- regions 134.
- the effusion aperture density of the second axial region 116 can be calculated as the average or mean of the densities of the first and second sub-regions 132 and 134.
- the support shell 30 of the first combustor wall 16 is located radially within the heat shield 32 of the first combustor wall 16.
- the heat shield 32 of the second combustor wall 20 is located radially within the support shell 30 of the second combustor wall 20.
- the heat shields 32 are respectively connected to the support shells 30 with a plurality of fasteners (e.g., heat shield studs and nuts).
- Each of the shell quench apertures 42 is fluidly coupled to a respective one of the shield quench apertures 94.
- one or more axial and/or circumferential impingement cavities are respectively defined between the support shell 30 and the heat shield 32.
- a first axial impingement cavity 136 extends between the support shell 30 and the panel 90 of the heat shield 32.
- Second and third axial impingement cavities 138 and 140 extend between the support shell 30 and the panel 92 of the heat shield 32.
- the first axial impingement cavity 136 respectively fluidly couples the first and second impingement apertures 52 and 54 with the first and second effusion apertures 104 and 106.
- first and/or second combustor walls 16 and 20 can be subject to axially and/or circumferentially varying combustion chamber 142 gas temperatures. Such varying
- the densities of the impingement and effusion apertures, for example, are relatively high adjacent regions of the combustion chamber 142 that have relatively high combustion chamber 142 gas temperatures. The densities of the
- impingement and effusion apertures are relatively low adjacent regions of the combustion chamber 142 that have relatively low combustion chamber 142 gas temperatures.
- the first and second combustor walls 16 and 20 can receive additional cooling air from the combustor plenum 144 in relatively hot regions of the combustion chamber 142 and less cooling air in relatively cool regions of the combustion chamber 142.
- the densities of the impingement and effusion apertures can be tailored such that the first and second combustor walls 16 and 20 are substantially isothermal during one or more modes of combustor 10 operation, which can reduce combustor wall material fatigue, etc.
- Cooling air flowing through the impingement apertures in the support shell 30 is subject to a cooling air first pressure drop between the combustor plenum surface 38 and the first impingement cavity surface 40.
- the magnitude of the first pressure drop is influenced by the number and/or diameter of the impingement apertures.
- Cooling air flowing through the effusion apertures in the heat shield 32 is subject to a cooling air second pressure drop between the second impingement cavity surface 86 and the combustion chamber surface 88.
- the magnitude of the second pressure drop is influenced by the number and/or diameter of the effusion apertures.
- the numbers and/or diameters of the impingement and effusion apertures are selected such that a ratio of the first pressure drop to the second pressure drop is between about two to one (2:1) and about nine to one (9:1).
- some or all of the axes 100 of the effusion apertures within a respective axial region of the heat shield 32 may be uniformly or non-uniformly aligned depending on, for example, (i) the flow and combustion temperatures of an adjacent region of the combustion chamber 142 and/or (ii) additional features (e.g., quench aperture, stud, etc.) included in the region.
- more than about seventy five percent (e.g., between about 80-100%) of the axes 100 of the third effusion apertures 108 in the third axial region 118 are aligned substantially perpendicular to the centerline 24 such that the cooling air flows into the combustion chamber 142 in a similar direction to the swirling combustion chamber 142 gas.
- the axes 100 of the fourth effusion apertures 96 in the fourth axial region 120 are arranged in various directions to cool the obstructed regions 128 surrounding the shield quench apertures 94.
- the axes 100 of the fourth effusion apertures 96 which are located downstream and adjacent to a respective one of the shield quench apertures 94 for example, are substantially tangent to a downstream side 146 of the shield quench aperture 94. In this manner, these fourth effusion apertures 96 can disturb stagnant flow regions within the combustion chamber 142; e.g., wake regions downstream of the shield quench apertures 94.
- the axes 100 of some of the first effusion apertures 104 are aligned substantially perpendicular to the centerline 24, while axes 100 of others of the first effusion apertures 104 are aligned substantially parallel to the centerline 24.
- suitable effusion (and impingement) aperture arrangements and alignments are disclosed in U.S. Patent No. 7,093,439, which is hereby incorporated by reference in its entirety.
- the impingement apertures 44 are offset from the effusion apertures 96. In this manner, the cooling air can impinge against and, thus, cool the second impingement cavity surface 86 before flowing into the effusion apertures 96.
- the effusion aperture density of one or more of the axial regions is between about one hundred and about three hundred effusion apertures per unit of combustion chamber surface 88.
- the effusion aperture density is relatively large where the angular offset between the effusion apertures and the combustion chamber surface 88 is relatively large (e.g., about thirty degrees).
- the effusion aperture density is relatively small where the angular offset between the effusion apertures and the combustion chamber surface 88 is relatively small (e.g., about fifteen degrees).
- one or more of the heat shields 32 includes a thermal barrier coating (TBC) applied to the combustion chamber surface 88.
- TBC thermal barrier coating
- the thermal barrier coating can include ceramic and/or any other suitable non-ceramic thermal barrier material.
- bosses surrounding the quench apertures may be interconnected and fluidly separate the cavity 138 into, for example, an axial forward cavity and an axial aft cavity.
Landscapes
- Engineering & Computer Science (AREA)
- Chemical & Material Sciences (AREA)
- Combustion & Propulsion (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Turbine Rotor Nozzle Sealing (AREA)
Abstract
Description
Claims
Applications Claiming Priority (2)
Application Number | Priority Date | Filing Date | Title |
---|---|---|---|
US13/531,132 US9052111B2 (en) | 2012-06-22 | 2012-06-22 | Turbine engine combustor wall with non-uniform distribution of effusion apertures |
PCT/US2013/047093 WO2013192540A1 (en) | 2012-06-22 | 2013-06-21 | Turbine engine combustor wall with non-uniform distribution of effusion apertures |
Publications (3)
Publication Number | Publication Date |
---|---|
EP2864707A1 true EP2864707A1 (en) | 2015-04-29 |
EP2864707A4 EP2864707A4 (en) | 2016-01-20 |
EP2864707B1 EP2864707B1 (en) | 2019-07-31 |
Family
ID=49769432
Family Applications (1)
Application Number | Title | Priority Date | Filing Date |
---|---|---|---|
EP13807403.4A Active EP2864707B1 (en) | 2012-06-22 | 2013-06-21 | Turbine engine combustor wall with non-uniform distribution of effusion apertures |
Country Status (3)
Country | Link |
---|---|
US (1) | US9052111B2 (en) |
EP (1) | EP2864707B1 (en) |
WO (1) | WO2013192540A1 (en) |
Families Citing this family (20)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
EP3039346B1 (en) * | 2013-08-30 | 2022-09-14 | Raytheon Technologies Corporation | Contoured dilution passages for a gas turbine engine combustor |
WO2015038232A1 (en) * | 2013-09-13 | 2015-03-19 | United Technologies Corporation | Sealed combustor liner panel for a gas turbine engine |
WO2015039074A1 (en) | 2013-09-16 | 2015-03-19 | United Technologies Corporation | Controlled variation of pressure drop through effusion cooling in a double walled combustor of a gas turbine engine |
EP2865850B1 (en) * | 2013-10-24 | 2018-01-03 | Ansaldo Energia Switzerland AG | Impingement cooling arrangement |
EP3084304B1 (en) | 2013-12-20 | 2020-08-26 | United Technologies Corporation | Cooling an aperture body of a combustor wall |
WO2015117137A1 (en) * | 2014-02-03 | 2015-08-06 | United Technologies Corporation | Film cooling a combustor wall of a turbine engine |
EP2921779B1 (en) * | 2014-03-18 | 2017-12-06 | Ansaldo Energia Switzerland AG | Combustion chamber with cooling sleeve |
US9851105B2 (en) | 2014-07-03 | 2017-12-26 | United Technologies Corporation | Self-cooled orifice structure |
GB201412460D0 (en) * | 2014-07-14 | 2014-08-27 | Rolls Royce Plc | An Annular Combustion Chamber Wall Arrangement |
EP2977679B1 (en) | 2014-07-22 | 2019-08-28 | United Technologies Corporation | Combustor wall for a gas turbine engine and method of acoustic dampening |
US10612781B2 (en) | 2014-11-07 | 2020-04-07 | United Technologies Corporation | Combustor wall aperture body with cooling circuit |
US10598382B2 (en) * | 2014-11-07 | 2020-03-24 | United Technologies Corporation | Impingement film-cooled floatwall with backside feature |
US20160258623A1 (en) * | 2015-03-05 | 2016-09-08 | United Technologies Corporation | Combustor and heat shield configurations for a gas turbine engine |
DE102015225825A1 (en) * | 2015-12-17 | 2017-06-22 | Rolls-Royce Deutschland Ltd & Co Kg | Gas turbine combustor with annular heat shield |
US20180266687A1 (en) * | 2017-03-16 | 2018-09-20 | General Electric Company | Reducing film scrubbing in a combustor |
US10753283B2 (en) * | 2017-03-20 | 2020-08-25 | Pratt & Whitney Canada Corp. | Combustor heat shield cooling hole arrangement |
US20190063322A1 (en) * | 2017-08-22 | 2019-02-28 | United Technologies Corporation | Hybrid floatwall cooling feature |
US10816202B2 (en) * | 2017-11-28 | 2020-10-27 | General Electric Company | Combustor liner for a gas turbine engine and an associated method thereof |
US11306919B2 (en) | 2018-10-19 | 2022-04-19 | Raytheon Technologies Corporation | Combustor panel cooling hole arrangement |
US11199326B2 (en) | 2019-12-20 | 2021-12-14 | Raytheon Technologies Corporation | Combustor panel |
Family Cites Families (38)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US4693667A (en) * | 1980-04-29 | 1987-09-15 | Teledyne Industries, Inc. | Turbine inlet nozzle with cooling means |
JPH0660740B2 (en) | 1985-04-05 | 1994-08-10 | 工業技術院長 | Gas turbine combustor |
US5687572A (en) | 1992-11-02 | 1997-11-18 | Alliedsignal Inc. | Thin wall combustor with backside impingement cooling |
US5758504A (en) | 1996-08-05 | 1998-06-02 | Solar Turbines Incorporated | Impingement/effusion cooled combustor liner |
GB2356924A (en) | 1999-12-01 | 2001-06-06 | Abb Alstom Power Uk Ltd | Cooling wall structure for combustor |
US6408629B1 (en) | 2000-10-03 | 2002-06-25 | General Electric Company | Combustor liner having preferentially angled cooling holes |
US6606861B2 (en) | 2001-02-26 | 2003-08-19 | United Technologies Corporation | Low emissions combustor for a gas turbine engine |
US6513331B1 (en) | 2001-08-21 | 2003-02-04 | General Electric Company | Preferential multihole combustor liner |
US7093439B2 (en) | 2002-05-16 | 2006-08-22 | United Technologies Corporation | Heat shield panels for use in a combustor for a gas turbine engine |
US6955053B1 (en) * | 2002-07-01 | 2005-10-18 | Hamilton Sundstrand Corporation | Pyrospin combuster |
US6964170B2 (en) | 2003-04-28 | 2005-11-15 | Pratt & Whitney Canada Corp. | Noise reducing combustor |
US7146815B2 (en) | 2003-07-31 | 2006-12-12 | United Technologies Corporation | Combustor |
US7093441B2 (en) | 2003-10-09 | 2006-08-22 | United Technologies Corporation | Gas turbine annular combustor having a first converging volume and a second converging volume, converging less gradually than the first converging volume |
US7036316B2 (en) | 2003-10-17 | 2006-05-02 | General Electric Company | Methods and apparatus for cooling turbine engine combustor exit temperatures |
US6982126B2 (en) | 2003-11-26 | 2006-01-03 | General Electric Company | Thermal barrier coating |
US6875529B1 (en) | 2003-12-30 | 2005-04-05 | General Electric Company | Thermal barrier coatings with protective outer layer for improved impact and erosion resistance |
US7291403B2 (en) | 2004-02-03 | 2007-11-06 | General Electric Company | Thermal barrier coating system |
US7326470B2 (en) | 2004-04-28 | 2008-02-05 | United Technologies Corporation | Thin 7YSZ, interfacial layer as cyclic durability (spallation) life enhancement for low conductivity TBCs |
US7464554B2 (en) | 2004-09-09 | 2008-12-16 | United Technologies Corporation | Gas turbine combustor heat shield panel or exhaust panel including a cooling device |
US7219498B2 (en) | 2004-09-10 | 2007-05-22 | Honeywell International, Inc. | Waffled impingement effusion method |
US7614235B2 (en) | 2005-03-01 | 2009-11-10 | United Technologies Corporation | Combustor cooling hole pattern |
US7597966B2 (en) | 2005-06-10 | 2009-10-06 | General Electric Company | Thermal barrier coating and process therefor |
US8084086B2 (en) | 2005-06-30 | 2011-12-27 | University Of Virginia Patent Foundation | Reliant thermal barrier coating system and related methods and apparatus of making the same |
US7416788B2 (en) | 2005-06-30 | 2008-08-26 | Honeywell International Inc. | Thermal barrier coating resistant to penetration by environmental contaminants |
US7856830B2 (en) * | 2006-05-26 | 2010-12-28 | Pratt & Whitney Canada Corp. | Noise reducing combustor |
JP2008093655A (en) | 2006-09-14 | 2008-04-24 | General Electric Co <Ge> | Method for preparing strain tolerant coating from green material |
US8021742B2 (en) | 2006-12-15 | 2011-09-20 | Siemens Energy, Inc. | Impact resistant thermal barrier coating system |
DE102007018061A1 (en) | 2007-04-17 | 2008-10-23 | Rolls-Royce Deutschland Ltd & Co Kg | Gas turbine combustion chamber wall |
US7905094B2 (en) * | 2007-09-28 | 2011-03-15 | Honeywell International Inc. | Combustor systems with liners having improved cooling hole patterns |
DE102008007870A1 (en) | 2008-02-06 | 2009-08-13 | Forschungszentrum Jülich GmbH | Thermal barrier coating system and process for its preparation |
US8056342B2 (en) | 2008-06-12 | 2011-11-15 | United Technologies Corporation | Hole pattern for gas turbine combustor |
US20100037620A1 (en) | 2008-08-15 | 2010-02-18 | General Electric Company, Schenectady | Impingement and effusion cooled combustor component |
US8161752B2 (en) * | 2008-11-20 | 2012-04-24 | Honeywell International Inc. | Combustors with inserts between dual wall liners |
GB0912715D0 (en) | 2009-07-22 | 2009-08-26 | Rolls Royce Plc | Cooling arrangement |
US9897320B2 (en) | 2009-07-30 | 2018-02-20 | Honeywell International Inc. | Effusion cooled dual wall gas turbine combustors |
US9416970B2 (en) | 2009-11-30 | 2016-08-16 | United Technologies Corporation | Combustor heat panel arrangement having holes offset from seams of a radially opposing heat panel |
US20110151132A1 (en) | 2009-12-21 | 2011-06-23 | Bangalore Nagaraj | Methods for Coating Articles Exposed to Hot and Harsh Environments |
US20110185739A1 (en) | 2010-01-29 | 2011-08-04 | Honeywell International Inc. | Gas turbine combustors with dual walled liners |
-
2012
- 2012-06-22 US US13/531,132 patent/US9052111B2/en active Active
-
2013
- 2013-06-21 WO PCT/US2013/047093 patent/WO2013192540A1/en active Application Filing
- 2013-06-21 EP EP13807403.4A patent/EP2864707B1/en active Active
Also Published As
Publication number | Publication date |
---|---|
US20130340437A1 (en) | 2013-12-26 |
EP2864707A4 (en) | 2016-01-20 |
WO2013192540A1 (en) | 2013-12-27 |
US9052111B2 (en) | 2015-06-09 |
EP2864707B1 (en) | 2019-07-31 |
Similar Documents
Publication | Publication Date | Title |
---|---|---|
EP2864707B1 (en) | Turbine engine combustor wall with non-uniform distribution of effusion apertures | |
CA2625330C (en) | Combustor liner with improved heat shield retention | |
EP2481983B1 (en) | Turbulated Aft-End liner assembly and cooling method for gas turbine combustor | |
JP4433529B2 (en) | Multi-hole membrane cooled combustor liner | |
JP5374031B2 (en) | Apparatus and gas turbine engine for making it possible to reduce NOx emissions in a turbine engine | |
US10317079B2 (en) | Cooling an aperture body of a combustor wall | |
US20100186415A1 (en) | Turbulated aft-end liner assembly and related cooling method | |
US20090120093A1 (en) | Turbulated aft-end liner assembly and cooling method | |
US10502422B2 (en) | Cooling a quench aperture body of a combustor wall | |
US20120304654A1 (en) | Combustion liner having turbulators | |
JP2010526274A (en) | Cooling holes for gas turbine combustor liners with non-uniform diameters therethrough | |
EP2475933A1 (en) | Fuel injector for use in a gas turbine engine | |
US10859271B2 (en) | Combustion chamber | |
JP2013513777A (en) | Turbine engine combustion chamber | |
EP2573464B1 (en) | Combustion sections of gas turbine engines with convection shield assemblies | |
EP3524885B1 (en) | Combustor panel standoff pin | |
EP3628927B1 (en) | Heat shield panel | |
US20100300107A1 (en) | Method and flow sleeve profile reduction to extend combustor liner life | |
US20130086915A1 (en) | Film cooled combustion liner assembly | |
US10697636B2 (en) | Cooling a combustor heat shield proximate a quench aperture | |
JPH04283315A (en) | Combustor liner | |
US20190249875A1 (en) | Liner for a Gas Turbine Engine Combustor | |
EP3447383B1 (en) | Hybrid floatwall cooling feature | |
US20140223912A1 (en) | Turbine engine combustion chamber | |
EP3604926B1 (en) | Heat shield panel for use in a gas turbine engine combustor |
Legal Events
Date | Code | Title | Description |
---|---|---|---|
PUAI | Public reference made under article 153(3) epc to a published international application that has entered the european phase |
Free format text: ORIGINAL CODE: 0009012 |
|
17P | Request for examination filed |
Effective date: 20150121 |
|
AK | Designated contracting states |
Kind code of ref document: A1 Designated state(s): AL AT BE BG CH CY CZ DE DK EE ES FI FR GB GR HR HU IE IS IT LI LT LU LV MC MK MT NL NO PL PT RO RS SE SI SK SM TR |
|
AX | Request for extension of the european patent |
Extension state: BA ME |
|
DAX | Request for extension of the european patent (deleted) | ||
RA4 | Supplementary search report drawn up and despatched (corrected) |
Effective date: 20151223 |
|
RIC1 | Information provided on ipc code assigned before grant |
Ipc: F02C 7/00 20060101ALI20151217BHEP Ipc: F23R 3/42 20060101AFI20151217BHEP Ipc: F02C 3/14 20060101ALI20151217BHEP |
|
RAP1 | Party data changed (applicant data changed or rights of an application transferred) |
Owner name: UNITED TECHNOLOGIES CORPORATION |
|
STAA | Information on the status of an ep patent application or granted ep patent |
Free format text: STATUS: EXAMINATION IS IN PROGRESS |
|
17Q | First examination report despatched |
Effective date: 20170620 |
|
GRAP | Despatch of communication of intention to grant a patent |
Free format text: ORIGINAL CODE: EPIDOSNIGR1 |
|
STAA | Information on the status of an ep patent application or granted ep patent |
Free format text: STATUS: GRANT OF PATENT IS INTENDED |
|
INTG | Intention to grant announced |
Effective date: 20190128 |
|
GRAS | Grant fee paid |
Free format text: ORIGINAL CODE: EPIDOSNIGR3 |
|
GRAA | (expected) grant |
Free format text: ORIGINAL CODE: 0009210 |
|
STAA | Information on the status of an ep patent application or granted ep patent |
Free format text: STATUS: THE PATENT HAS BEEN GRANTED |
|
AK | Designated contracting states |
Kind code of ref document: B1 Designated state(s): AL AT BE BG CH CY CZ DE DK EE ES FI FR GB GR HR HU IE IS IT LI LT LU LV MC MK MT NL NO PL PT RO RS SE SI SK SM TR |
|
REG | Reference to a national code |
Ref country code: CH Ref legal event code: EP Ref country code: GB Ref legal event code: FG4D |
|
REG | Reference to a national code |
Ref country code: AT Ref legal event code: REF Ref document number: 1161289 Country of ref document: AT Kind code of ref document: T Effective date: 20190815 |
|
REG | Reference to a national code |
Ref country code: IE Ref legal event code: FG4D |
|
REG | Reference to a national code |
Ref country code: DE Ref legal event code: R096 Ref document number: 602013058559 Country of ref document: DE |
|
REG | Reference to a national code |
Ref country code: NL Ref legal event code: MP Effective date: 20190731 |
|
REG | Reference to a national code |
Ref country code: LT Ref legal event code: MG4D |
|
REG | Reference to a national code |
Ref country code: AT Ref legal event code: MK05 Ref document number: 1161289 Country of ref document: AT Kind code of ref document: T Effective date: 20190731 |
|
PG25 | Lapsed in a contracting state [announced via postgrant information from national office to epo] |
Ref country code: NL Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT Effective date: 20190731 Ref country code: BG Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT Effective date: 20191031 Ref country code: SE Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT Effective date: 20190731 Ref country code: LT Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT Effective date: 20190731 Ref country code: NO Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT Effective date: 20191031 Ref country code: PT Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT Effective date: 20191202 Ref country code: AT Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT Effective date: 20190731 Ref country code: FI Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT Effective date: 20190731 Ref country code: HR Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT Effective date: 20190731 |
|
PG25 | Lapsed in a contracting state [announced via postgrant information from national office to epo] |
Ref country code: LV Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT Effective date: 20190731 Ref country code: AL Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT Effective date: 20190731 Ref country code: RS Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT Effective date: 20190731 Ref country code: IS Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT Effective date: 20191130 Ref country code: GR Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT Effective date: 20191101 Ref country code: ES Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT Effective date: 20190731 |
|
PG25 | Lapsed in a contracting state [announced via postgrant information from national office to epo] |
Ref country code: TR Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT Effective date: 20190731 |
|
PG25 | Lapsed in a contracting state [announced via postgrant information from national office to epo] |
Ref country code: DK Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT Effective date: 20190731 Ref country code: IT Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT Effective date: 20190731 Ref country code: PL Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT Effective date: 20190731 Ref country code: EE Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT Effective date: 20190731 Ref country code: RO Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT Effective date: 20190731 |
|
PG25 | Lapsed in a contracting state [announced via postgrant information from national office to epo] |
Ref country code: IS Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT Effective date: 20200224 Ref country code: CZ Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT Effective date: 20190731 Ref country code: SK Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT Effective date: 20190731 Ref country code: SM Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT Effective date: 20190731 |
|
REG | Reference to a national code |
Ref country code: DE Ref legal event code: R097 Ref document number: 602013058559 Country of ref document: DE |
|
PLBE | No opposition filed within time limit |
Free format text: ORIGINAL CODE: 0009261 |
|
STAA | Information on the status of an ep patent application or granted ep patent |
Free format text: STATUS: NO OPPOSITION FILED WITHIN TIME LIMIT |
|
PG2D | Information on lapse in contracting state deleted |
Ref country code: IS |
|
PG25 | Lapsed in a contracting state [announced via postgrant information from national office to epo] |
Ref country code: IS Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT Effective date: 20191030 |
|
26N | No opposition filed |
Effective date: 20200603 |
|
PG25 | Lapsed in a contracting state [announced via postgrant information from national office to epo] |
Ref country code: SI Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT Effective date: 20190731 |
|
PG25 | Lapsed in a contracting state [announced via postgrant information from national office to epo] |
Ref country code: MC Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT Effective date: 20190731 |
|
REG | Reference to a national code |
Ref country code: CH Ref legal event code: PL |
|
PG25 | Lapsed in a contracting state [announced via postgrant information from national office to epo] |
Ref country code: LU Free format text: LAPSE BECAUSE OF NON-PAYMENT OF DUE FEES Effective date: 20200621 |
|
REG | Reference to a national code |
Ref country code: BE Ref legal event code: MM Effective date: 20200630 |
|
PG25 | Lapsed in a contracting state [announced via postgrant information from national office to epo] |
Ref country code: IE Free format text: LAPSE BECAUSE OF NON-PAYMENT OF DUE FEES Effective date: 20200621 Ref country code: LI Free format text: LAPSE BECAUSE OF NON-PAYMENT OF DUE FEES Effective date: 20200630 Ref country code: CH Free format text: LAPSE BECAUSE OF NON-PAYMENT OF DUE FEES Effective date: 20200630 |
|
PG25 | Lapsed in a contracting state [announced via postgrant information from national office to epo] |
Ref country code: BE Free format text: LAPSE BECAUSE OF NON-PAYMENT OF DUE FEES Effective date: 20200630 |
|
PG25 | Lapsed in a contracting state [announced via postgrant information from national office to epo] |
Ref country code: MT Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT Effective date: 20190731 Ref country code: CY Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT Effective date: 20190731 |
|
PG25 | Lapsed in a contracting state [announced via postgrant information from national office to epo] |
Ref country code: MK Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT Effective date: 20190731 |
|
REG | Reference to a national code |
Ref country code: DE Ref legal event code: R081 Ref document number: 602013058559 Country of ref document: DE Owner name: RAYTHEON TECHNOLOGIES CORPORATION (N.D.GES.D.S, US Free format text: FORMER OWNER: UNITED TECHNOLOGIES CORPORATION, FARMINGTON, CONN., US |
|
P01 | Opt-out of the competence of the unified patent court (upc) registered |
Effective date: 20230520 |
|
PGFP | Annual fee paid to national office [announced via postgrant information from national office to epo] |
Ref country code: GB Payment date: 20240521 Year of fee payment: 12 |
|
PGFP | Annual fee paid to national office [announced via postgrant information from national office to epo] |
Ref country code: DE Payment date: 20240521 Year of fee payment: 12 |
|
PGFP | Annual fee paid to national office [announced via postgrant information from national office to epo] |
Ref country code: FR Payment date: 20240522 Year of fee payment: 12 |